U.S. patent number 11,015,475 [Application Number 16/233,964] was granted by the patent office on 2021-05-25 for passive blade tip clearance control system for gas turbine engine.
This patent grant is currently assigned to Rolls-Royce Corporation, Rolls-Royce North American Technologies Inc.. The grantee listed for this patent is Rolls-Royce Corporation, Rolls-Royce North American Technologies Inc.. Invention is credited to Ryan C. Humes, Behram Kapadia, Jack Moody, Brandon R. Snyder.
United States Patent |
11,015,475 |
Kapadia , et al. |
May 25, 2021 |
Passive blade tip clearance control system for gas turbine
engine
Abstract
The present disclosure relates to a gas turbine engine including
a turbine wheel mounted for rotation about a central axis and a
turbine shroud ring mounted radially outward from the turbine
wheel. The turbine wheel includes a plurality of blades that are
spaced apart radially from the turbine shroud ring to establish a
blade tip clearance gap. The gas turbine engine further includes a
blade tip clearance control system that passively controls the size
of the clearance gap based on engine operation.
Inventors: |
Kapadia; Behram (McCordsville,
IN), Moody; Jack (Indianapolis, IN), Humes; Ryan C.
(Indianapolis, IN), Snyder; Brandon R. (Greenwood, IN) |
Applicant: |
Name |
City |
State |
Country |
Type |
Rolls-Royce Corporation
Rolls-Royce North American Technologies Inc. |
Indianapolis
Indianapolis |
IN
IN |
US
US |
|
|
Assignee: |
Rolls-Royce Corporation
(Indianapolis, IN)
Rolls-Royce North American Technologies Inc. (Indianapolis,
IN)
|
Family
ID: |
1000005574389 |
Appl.
No.: |
16/233,964 |
Filed: |
December 27, 2018 |
Prior Publication Data
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|
|
|
Document
Identifier |
Publication Date |
|
US 20200208533 A1 |
Jul 2, 2020 |
|
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F02K
3/06 (20130101); F01D 25/12 (20130101); F01D
11/18 (20130101); F02C 9/18 (20130101); F05D
2220/323 (20130101); F05D 2240/11 (20130101); F05D
2220/3212 (20130101); F05D 2270/42 (20130101); F01D
19/02 (20130101); F05D 2260/606 (20130101) |
Current International
Class: |
F01D
11/18 (20060101); F01D 25/12 (20060101); F02C
9/18 (20060101); F02K 3/06 (20060101); F01D
19/02 (20060101) |
References Cited
[Referenced By]
U.S. Patent Documents
Foreign Patent Documents
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1004759 |
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May 2000 |
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EP |
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1630385 |
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Mar 2006 |
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EP |
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3176382 |
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Jun 2017 |
|
EP |
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2004097181 |
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Nov 2004 |
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WO |
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2015094990 |
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Jun 2015 |
|
WO |
|
Other References
European Search Report for Application No. 19211684.6-1004, dated
Mar. 9, 2020, 7 pages. cited by applicant.
|
Primary Examiner: Seabe; Justin D
Attorney, Agent or Firm: Barnes & Thornburg LLP
Claims
What is claimed is:
1. A gas turbine engine comprising a compressor configured to
pressurize air moving along a primary gas path of the gas turbine
engine, a combustor fluidly coupled to the compressor to receive
pressurized air discharged from the compressor and configured to
ignite fuel mixed with the pressurized air, and a turbine including
(i) a high-pressure section fluidly coupled to the combustor to
receive combustion gases generated by fuel burned in the combustor
and (ii) a low-pressure section fluidly coupled to receive the
combustion gasses exiting the high-pressure section, wherein the
high-pressure section includes a turbine wheel mounted for rotation
about a central reference axis, a variable-diameter turbine shroud
ring that extends around the turbine wheel, and a passive blade-tip
clearance control system including a shroud-ring support coupled to
the variable-diameter turbine shroud ring that is configured to
drive motion of the turbine shroud ring radially inward or outward
based on temperature of the shroud-ring support and defining at
least in part a cavity located radially outward of the
variable-diameter turbine shroud ring, wherein the cavity is
fluidly coupled to a bleed-air passageway that extends from the
compressor to the cavity without interruption from a valve and a
cooling-air passageway that extends from the cavity to the low
pressure section such that pressurized bleed air from the
compressor is conducted to the cavity of the passive blade tip
clearance control system so that the temperature and motion of the
shroud-ring support is controlled based on the operating conditions
of the engine without active control of the pressurized bleed air
provided to the cavity, wherein the passive blade tip clearance
control system further includes an outer case and the shroud-ring
support is mounted radially-inward of the outer case to define the
cavity radially between an inner surface of the outer case and an
outer surface of the shroud-ring support so that pressurized bleed
air from the compressor passes over the outer surface of the
shroud-ring support, wherein the shroud-ring support includes a
panel that is coupled to the turbine shroud ring and a flange
coupled to an axially-forward end of the panel, the flange coupled
to the outer case and having a U-shape when viewed
circumferentially so that the flange is configured to flex as the
shroud-ring support moves radially inward and outward relative to
the outer case as the pressurized bleed air drives movement of the
shroud-ring support, wherein the outer case further includes an
outer panel that is concentric with the panel of the shroud-ring
support and defines the inner surface of the outer case, an annular
duct that extends circumferentially around the reference axis and
defines a manifold in fluid communication with the bleed-air
passageway, and an inner panel that extends axially forward from
the annular duct at an outlet of the manifold and is located
radially inward of the outer panel and radially outward of the
shroud-ring support to define a gap radially therebetween that is
configured to accelerate a flow of the pressurized bleed air from
the manifold axially forward over the outer surface of the
shroud-ring support toward the flange.
2. The gas turbine engine of claim 1, wherein the passive blade tip
clearance control system further includes an inlet conduit coupled
to the outer case and opening into the cavity and an outlet, the
inlet configured to conduct the bleed air from the compressor into
the cavity and the outlet configured to conduct the bleed air from
the cavity to the low pressure section of the turbine.
3. The gas turbine engine of claim 2, wherein the passive blade-tip
clearance control system is configured to heat the shroud-ring
support during start-up conditions of the gas turbine engine and is
configured to cool the shroud-ring support during cruise
conditions.
4. The gas turbine engine of claim 1, wherein the cavity formed
between the outer case and the shroud-ring support is sealed off
from a gas path of the high pressure section of the turbine such
that the temperature of gases within the cavity controls the gap
while allowing for pressure within the cavity to be less than
pressure within the primary gas path of the high pressure turbine
section.
5. The gas turbine engine of claim 1, wherein the passive blade tip
clearance control system includes a plurality of inlet conduits
fluidly coupled to the manifold and spaced apart circumferentially
around the reference axis and a plurality of outlets spaced apart
circumferentially around the reference axis that extend through the
manifold and are offset from each inlet conduit.
6. The gas turbine engine of claim 1, wherein the high pressure
section of the turbine includes a first turbine blade stage, a
second turbine blade stage axially aft of the first turbine blade
stage, and a vane stage axially between the first and second
turbine blade stages, and the passive blade tip clearance control
system is configured to control a gap radially between second
turbine blade stage and the turbine shroud ring.
7. The gas turbine engine of claim 6, wherein the outer panel is
spaced apart from the central reference axis a first distance, and
the inner panel is spaced apart from the central reference axis a
second distance that is less than the first distance.
8. The gas turbine engine of claim 7, wherein the inner panel is
positioned radially outward of the second turbine blade stage such
that the cavity is narrowed outward of the second turbine blade
stage.
9. The gas turbine engine of claim 7, wherein the inner panel is
adjustable axially to target additional turbine blade stages
included in the high pressure section of the turbine.
10. The gas turbine engine of claim 6, wherein the shroud-ring
support includes a plurality of turbulators coupled to the outer
surface of the shroud-ring support within the cavity radially
outward of the second turbine blade stage to increase heat transfer
between the bleed air and the shroud-ring support directly outward
of the second turbine blade stage.
11. The gas turbine engine of claim 1, wherein the flange includes
a radially inner flex-section coupled to the panel that extends
axially forward from the pane, a radially outer flex-section
coupled to the radially inner flex-section that extends axially aft
from the radially inner flex-section, and a mount section that
extends radially outward from the radially outer flex-section and
couples to the outer case to mount the shroud-ring support to and
the turbine shroud ring to the outer case, wherein the radially
inner flex-section and the radially outer flex-section are arranged
at an angle relative to one another to provide the U-shape of the
flange.
12. A high pressure turbine section for use in a gas turbine
engine, the turbine section comprising a turbine wheel mounted for
rotation about a central reference axis, a plurality of blades that
extend radially outward from the turbine wheel to interact with
gases moving through a primary gas path of the turbine section, a
variable-diameter turbine shroud ring that extends around the
turbine wheel to define a radially-outer boundary of the primary
gas path, and a passive blade-tip clearance control system
configured to drive motion of the turbine shroud ring radially
inward and outward relative to the central reference axis to
control size of a gap radially between the turbine wheel and the
variable-diameter turbine shroud ring, the passive blade-tip
clearance control system including an outer case, a shroud-ring
support mounted radially-inward of the outer case to define a
cavity radially therebetween, and an inlet conduit coupled to the
outer case, wherein the cavity formed between a radially-inwardly
facing surface of the outer case and a radially-outwardly facing
surface of the shroud-ring support is sealed off from the primary
gas path within the high pressure turbine section, wherein the
shroud-ring support includes a panel that is coupled to the turbine
shroud ring and defines the radially-outward facing surface of the
shroud-ring support and a flange coupled to an axially-forward end
of the panel, the flange coupled to the outer case and having a
U-shape when viewed circumferentially, and wherein the inlet
conduit defines a bleed air passageway that opens into the cavity
and is configured to conduct bleed air into the cavity across the
radially-outwardly facing surface of the shroud-ring support to
cause the panel to move radially inward or outward and the flange
is configured to flex as the shroud-ring support moves radially
inward or outward relative to the outer case as the bleed air
drives movement of the shroud-ring support.
13. The high pressure turbine section of claim 12, wherein the
outer case includes an outer panel spaced apart from the central
reference axis a first distance, and an inner panel spaced apart
from the central reference axis a second distance that is less than
the first distance.
14. The high pressure turbine section of claim 13, wherein the high
pressure section of the turbine includes a first turbine blade
stage, a second turbine blade stage axially aft of the first
turbine blade stage, and a vane stage axially between the first and
second turbine blade stages, and wherein the inner panel is
positioned radially outward of the second turbine blade stage such
that the cavity is narrowed outboard of the second turbine blade
stage.
15. The high pressure turbine section of claim 13, wherein the
inner panel is adjustable axially to target additional turbine
blade stages included in the high pressure section of the
turbine.
16. The high pressure turbine section of claim 12, wherein the
shroud-ring support includes a plurality of turbulators coupled to
the radially-outwardly facing surface of the shroud-ring support
within the cavity.
17. The high pressure turbine section of claim 12, wherein the
panel of the shroud-ring support including an axially-aft end
spaced apart axially aft of the axially-forward end of the panel
that is free for radial movement relative to the outer case.
18. The high pressure turbine section of claim 12, wherein the
flange includes a radially inner flex-section coupled to the panel
that extends axially forward from the pane, a radially outer
flex-section coupled to the radially inner flex-section that
extends axially aft from the radially inner flex-section, and a
mount section that extends radially outward from the radially outer
flex-section and couples to the outer case to mount the shroud-ring
support to and the turbine shroud ring to the outer case, wherein
the radially inner flex-section and the radially outer flex-section
are arranged at an angle relative to one another to provide the
U-shape of the flange.
19. The high pressure turbine section of claim 18, wherein the
angle between the radially inner flex-section and the radially
outer flex section increases as the shroud-ring support moves
radially inward in response to the bleed air flowing across the
radially-outer surface of the shroud-ring support, and wherein the
angle between the radially inner flex-section and radially the
outer flex section decreases as the shroud-ring support moves
radially outward in response to the bleed air flowing across the
radially-outer surface of the shroud-ring support.
Description
FIELD OF THE DISCLOSURE
The present disclosure relates generally to a gas turbine engine
including a blade tip clearance control system. More particularly,
the present disclosure relates to a passive blade tip clearance
control system.
BACKGROUND
Gas turbine engines are used to power aircraft, watercraft, power
generators, and the like. Gas turbine engines typically include a
compressor, a combustor, and a turbine. The compressor compresses
air drawn into the engine and delivers high pressure air to the
combustor. In the combustor, fuel is mixed with the high pressure
air and is ignited. Products of the combustion reaction in the
combustor are directed into the turbine where work is extracted to
drive the compressor and, sometimes, an output shaft. Left-over
products of the combustion are exhausted out of the turbine and may
provide thrust in some applications.
Products of the combustion reaction directed into the turbine flow
over airfoils included in rotating blades of the turbine. A blade
track or other structure arranged radially outward of the rotating
blades may block combustion products from passing over the blades
without causing the blades to rotate, thereby contributing to lost
performance within the gas turbine engine. Excessive contact
between the rotating blades and the blade track during engine
operation may cause degradation of the blades. Excessive clearance
between the rotating blades and the blade track may cause
unacceptable efficiencies of the gas turbine engine. In view of the
above considerations, managing clearance between the blade track
and the rotating blades remains an area of interest.
SUMMARY
The present disclosure may comprise one or more of the following
features and combinations thereof.
A gas turbine engine, in accordance with the present disclosure,
includes a compressor, a combustor, and a turbine. The compressor
is configured to pressurize air moving along a primary gas path of
the gas turbine engine. The combustor is fluidly coupled to the
compressor to receive pressurized air discharged from the
compressor and configured to ignite fuel mixed with the pressurized
air. The turbine includes (i) a high-pressure section fluidly
coupled to the combustor to receive combustion gases generated by
fuel burned in the combustor and (ii) a low-pressure section
fluidly coupled to receive the combustion gasses exiting the
high-pressure section.
In illustrative embodiments, the high-pressure section includes a
turbine wheel mounted for rotation about a central reference axis,
a variable-diameter turbine shroud ring that extends around the
turbine wheel, and a passive blade-tip clearance control system.
The passive blade tip clearance control system includes a
shroud-ring support coupled to the variable-diameter turbine shroud
ring that is configured to drive motion of the turbine shroud ring
radially inward or outward based on temperature of the shroud-ring
support and defining at least in part a cavity located radially
outward of the variable-diameter turbine shroud ring.
In illustrative embodiments, the cavity is fluidly coupled to a
bleed-air passageway that extends from the compressor to the cavity
without interruption from a valve and a cooling-air passageway that
extends from the cavity to the low pressure section such that
pressurized bleed air from the compressor is conducted to the
cavity of the passive blade tip clearance control system so that
the temperature and motion of the shroud-ring support is controlled
based on the operating conditions of the engine without active
control of the pressurized bleed air provided to the cavity.
In illustrative embodiments, the passive blade tip clearance
control system further includes an outer case and the shroud-ring
support is provided by an inner case mounted radially-inward of the
outer case to define the cavity radially therebetween.
In illustrative embodiments, the passive blade tip clearance
control system further includes an inlet conduit coupled to the
outer case and opening into the cavity and an outlet, the inlet
configured to conduct the bleed air from the compressor into the
cavity and the outlet configured to conduct the bleed air from the
cavity to the low pressure section of the turbine.
In illustrative embodiments, the passive blade-tip clearance
control system is configured to heat the inner case during start-up
conditions of the gas turbine engine and is configured to cool the
inner case during cruise conditions.
In illustrative embodiments, the cavity formed between the outer
case and the inner case is sealed off from a gas path of the high
pressure section of the turbine such that the temperature of gases
within the cavity controls the gap while allowing for pressure
within the cavity to be less than pressure within the primary gas
path of the high pressure turbine section.
In illustrative embodiments, the outer case includes an annular
duct that extends circumferentially around the reference axis and
defines a manifold and the inlet is fluidly coupled to the
manifold.
In illustrative embodiments, the passive blade tip clearance
control system includes a plurality of inlet conduits fluidly
coupled to the manifold and spaced apart circumferentially around
the reference axis and a plurality of outlets spaced apart
circumferentially around the reference axis that extend through the
manifold and are offset from each inlet conduit.
In illustrative embodiments, the high pressure section of the
turbine includes a first turbine blade stage, a second turbine
blade stage axially aft of the first turbine blade stage, and a
vane stage axially between the first and second turbine blade
stages. The passive blade tip clearance control system is
configured to control a gap radially between second turbine blade
stage and the turbine shroud ring.
In illustrative embodiments, the outer case includes an outer panel
spaced apart from the central reference axis a first distance and
an inner panel spaced apart from the central reference axis a
second distance that is less than the first distance.
In illustrative embodiments, the inner panel is positioned radially
outward of the second turbine blade stage such that the cavity is
narrowed outward of the second turbine blade stage.
In illustrative embodiments, the outer panel is spaced apart
axially from the inner panel.
In illustrative embodiments, the inner panel is adjustable axially
to target additional turbine blade stages included in the high
pressure section of the turbine.
In illustrative embodiments, the inner case includes a plurality of
turbulators coupled to an upper surface of the inner case within
the cavity radially outward of the second turbine blade stage to
increase heat transfer between the bleed air and the inner case
directly outward of the second turbine blade stage.
In illustrative embodiments, the inner case includes a panel that
is coupled to the turbine shroud ring and a flange coupled to an
axially-forward end of the panel, the flange coupled to the outer
case and having a U-shape when viewed circumferentially so that the
flange is configured to flex as the inner case moves radially
inward and outward relative to the outer case.
According to another aspect of the present disclosure, a high
pressure turbine section for use in a gas turbine engine includes a
turbine wheel mounted for rotation about a central reference axis,
a plurality of blades that extend radially outward from the turbine
wheel to interact with gases moving through a primary gas path of
the turbine section, a variable-diameter turbine shroud ring that
extends around the turbine wheel to define a radially-outer
boundary of the primary gas path, and a passive blade-tip clearance
control system.
In illustrative embodiments, the passive blade tip clearance
control system is configured to drive motion of the turbine shroud
ring radially inward and outward relative to the central reference
axis to control size of a gap radially between the turbine wheel
and the variable-diameter turbine shroud ring, the passive
blade-tip clearance control system including an outer case and an
inner case mounted radially-inward of the outer case to define a
cavity radially therebetween. The cavity formed between the outer
case and the inner case is sealed off from the primary gas path
within the high pressure turbine section.
These and other features of the present disclosure will become more
apparent from the following description of the illustrative
embodiments.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 is a perspective view of a gas turbine engine including a
passive blade tip clearance control system integrated into a
turbine of the engine as shown in FIG. 2, the passive blade tip
clearance system being configured to conduct pressurized bleed air
from a compressor of the engine to a high-pressure section of the
turbine to control blade tip clearance in the high-pressure section
without interruption from valves or other active control
devices;
FIG. 2 is a cross sectional view of the high-pressure section of
the turbine showing that the passive blade tip clearance control
system includes an outer case and an inner case that cooperate to
define a cavity, and showing that the cavity is configured to
receive the bleed air from the compressor to drive movement of the
inner case radially inward or outward depending on operating
conditions of the engine to control blade tip clearance;
FIG. 3 is a perspective view of the passive blade tip clearance
control system of FIGS. 1 and 2 showing that the passive blade tip
clearance control system includes a plurality of inlets spaced
circumferentially around a central reference axis and a plurality
of outlets spaced circumferentially around the reference axis;
FIG. 4 is a cross sectional view of a second passive blade tip
clearance control system, in accordance with the present
disclosure, the passive blade tip clearance control system
including an outer case and an inner case that define a cavity
therebetween, and showing an inlet opening into the cavity through
the outer case to conduct bleed air into the cavity and an outlet
downstream of the inlet where the cavity is narrowed by the outer
case to accelerate the bleed air through the outlet; and
FIG. 5 is a cross sectional view of a third passive blade tip
clearance control system, in accordance with the present
disclosure, the passive blade tip clearance control system
including an outer case and an inner case that define a cavity
therebetween, and showing an inlet opening into the cavity through
the outer case to conduct bleed air into the cavity and an outlet
downstream of the inlet where the cavity is narrowed by a separate
turbine shroud component to accelerate the bleed air through the
outlet.
DETAILED DESCRIPTION
For the purposes of promoting an understanding of the principles of
the disclosure, reference will now be made to a number of
illustrative embodiments illustrated in the drawings and specific
language will be used to describe the same.
An illustrative aerospace gas turbine engine 10 includes a fan 12,
a compressor 14, a combustor 16, and a turbine 18 as shown in FIG.
1. The fan 12 is driven by the turbine 18 and provides thrust for
propelling an air vehicle. The compressor 14 compresses and
delivers air to the combustor 16. The combustor 16 mixes fuel with
the compressed air received from the compressor 14 and ignites the
fuel. The hot, high-pressure products of the combustion reaction in
the combustor 16 are directed into the turbine 18 to cause the
turbine 18 to rotate about a central reference axis 11 and drive
the compressor 14 and the fan 12.
The turbine 18 includes a high-pressure section 20 fluidly coupled
to the combustor 16 to receive combustion gases generated by fuel
burned in the combustor and a low-pressure section 22 fluidly
coupled to receive the combustion gasses exiting the high-pressure
section 20. In other embodiments, the turbine 18 may further
include one or more intermediate sections between the high pressure
section 20 and the low pressure section 22.
The high-pressure section 20 includes a turbine wheel 24 mounted
for rotation about the central reference axis 11, a
variable-diameter turbine shroud ring 26 that extends around the
turbine wheel 24, and a plurality of blades 30 mounted to the
turbine wheel 24 as shown in FIGS. 1 and 2. The blades 20 are
configured to interact with the hot gases in the gas path to cause
rotation of the turbine wheel 24. The turbine shroud ring 26 is
coupled to a shroud ring support 34 radially outward from the
blades 30 to provide a radially outer boundary of the gas path.
Each of the blades 30 is spaced apart from the turbine shroud ring
26 by a clearance gap 36 that is defined between a tip of each
blade 30 and the turbine shroud ring 26.
During operation of the gas turbine engine, the hot gases may cause
various components within the turbine 18 to expand in response to
the higher temperatures caused by the combustion of the pressurized
air. One such component that may undergo thermal expansion is the
shroud-ring support 34. For example, the shroud ring support 34 may
have a first diameter under start-up conditions and a second
diameter under operating conditions that is greater than the first
diameter due to the increased temperature of the engine after a
period of time of operation. Similarly, the turbine wheel 24 may
have a first diameter under start-up conditions and a second
diameter under operating conditions that is greater than the first
diameter due to the increased temperature of the engine after a
period of time of operation. However, the turbine wheel 24 changes
from the first diameter to the second diameter in a shorter period
of time than the shroud-ring support 34.
In order to control the gap 36 while the turbine wheel 24 and the
shroud-ring support change diameter relative to one another, the
high pressure section 20 further includes a passive blade-tip
clearance control system 40 that defines a cavity 50 radially
outward from the turbine wheel 24 of the high pressure section 20
as shown in FIGS. 1 and 2. The cavity 50 is sealed from the gas
path of the turbine 18 such that the temperature of gases within
the cavity 50 controls the gap 36 while allowing for pressure
within the cavity 50 to be less than pressure within the primary
gas path of the high pressure turbine section 20.
During start-up of the gas turbine engine 10, the passive blade tip
clearance control system 40 supplies pressurized air having a first
temperature to the cavity 50. At this point in time, the
temperature of the pressurized air is greater than the temperature
of the shroud-ring support 34 to heat the shroud-ring support 34.
Heating the shroud-ring support 34 during start-up opens the gap 36
to accommodate the faster change in diameter of the turbine wheel
24. In other words, heating the shroud-ring support 34 during
start-up allows the thermal growth of the shroud-ring support 34 to
more closely match the thermal growth of the turbine wheel 24.
During operating conditions of the engine 10, such as cruise
conditions, the passive blade tip clearance control system 40
supplies the same pressurized air to the cavity 50. However, under
operating conditions, after a period of time, the temperature of
the shroud-ring support 34 becomes greater than the temperature of
the pressurized air. At this point in time the pressurized air
cools the shroud-ring support 34 to close the gap 36 and improve
efficiencies of the engine 10.
The passive blade tip clearance control system 40 includes an inlet
conduit 42, an outer case 44 and the shroud-ring support 34, or
inner case 34. The outer case 44 and the inner case 34 define the
cavity 50. The inlet conduit 42 defines a bleed-air passageway 46
that is fluidly coupled to the compressor 14 and the cavity 50
without interruption from a valve or any other active blade tip
clearance control device. A cooling-air passageway 47 extends from
the cavity 50 to the low pressure section 22 of the turbine 18. In
this way, the blade tip clearance control system 40 conducts
pressurized air from the compressor 14 into the cavity 50 to
control blade tip clearance in the high pressure section 20 and
then reuses the air downstream of the cavity 50 in the low pressure
section 22. Furthermore, the temperature and motion of the
shroud-ring support 34 is controlled based on the operating
conditions of the engine without active control of the pressurized
bleed air provided to the cavity.
The outer case 44 includes an outer panel 52, an inner panel 54
spaced apart radially from the outer panel 52, and an annular duct
56 as shown in FIGS. 2 and 3. The outer panel 52 defines a
radially-outer boundary of the cavity 50. The inner panel 54
extends axially forward from the annular duct 56 within the cavity
50 such that the inner panel 54 is spaced apart radially between
both the outer panel 52 and the shroud-ring support 34. The annular
duct 56 defines a manifold 58 that extends circumferentially around
the central reference axis 11 as shown in FIG. 3. The bleed-air
passageway 46 has an outlet 48 that opens into the manifold 56 to
deliver the pressurized air into the manifold 58.
The shroud-ring support 34, or inner case 34, includes a panel 60,
a flange 62 on an axially forward end of the panel 60, and a
plurality of mounts 63 that extends radially downward from the
panel 60 as shown in FIG. 2. The panel 60 defines a radially-inner
boundary of the cavity 50. The flange 62 has a generally U-shape
cross section when viewed circumferentially and extends radially
outward into engagement with the outer case 44 to provide the
cavity 50. The plurality of mounts 63 interface with corresponding
mounts 65 coupled to the turbine shroud ring 26 to mount the
turbine shroud ring 26 to the shroud-ring support 34. In the
illustrative embodiment, the mounts 63, 65 are generally L-shaped
when viewed circumferentially, however, in other embodiments, the
mounts may have any suitable shape. Additionally, one or more
fasteners may be used to mount the turbine shroud ring 26 to the
shroud-ring support 34.
The annular duct 56 is a tubular member to define the manifold 58
as shown in FIG. 2. The annular duct 56 is formed to include a gap
64 that extends circumferentially around the reference axis 11 with
the annular duct 56. The gap 64 acts as an outlet to the manifold
and opens into the cavity 50. The inner panel 54 is coupled to the
annular duct and extends axially forward from the annular duct 56
into the cavity to provide the gap 64 along an axial length 66 of
the inner panel 54 into the cavity 50. The inner panel 54
cooperates with the annular duct 56 to narrow the cavity 50 along
the length 66. As the cavity 50 is narrowed, the pressurized air is
accelerated in that area to encourage heat transfer between the
pressurized air and a portion of the shroud ring support 34
corresponding to the length 66 to control blade tip clearance in
that area.
The inner panel 54 may be extended or retracted axially to adjust
the length 66 and, thus, the portion of the shroud-ring support 34
that is controlled. Additionally, the length or length 66 of the
inner panel 54 may be increased or decreased based on the operating
conditions of the particular engine in which the passive blade tip
clearance control system 40 is included. In one embodiment, an
amount of radial movement of the shroud-ring support 34 is related
to the length 66 of the inner panel 54. For example, if more blade
tip clearance is required in a particular engine, the length 66 of
the inner panel 54 may be increased to provide greater heat
transfer across a larger portion of the shroud ring support 34 and,
thus, a larger change in diameter of the shroud ring support 34
relative to the reference axis 11. However, it should be noted that
any suitable length 66 may be used depending on the particular
application.
The shroud-ring support 34 may further include a plurality of
turbulators 90 formed on the shroud ring support 34 as shown in
FIG. 2. The turbulators 90 are positioned in the gap 64 within the
length 66 of the inner panel 54 to increase the surface area of the
shroud ring support 34 in that area. The turbulators 90 increase
heat transfer from the pressurized air to the shroud-ring support
34 to help drive the shroud ring support 34 radially inward and
outward. However, in other embodiments, the turbulators 90 may
extend along the shroud-ring support 34 a greater or lesser
distance than the length 66 of the inner panel 54.
The flange 62 is configured to flex to allow the shroud-ring
support 34 to move radially inward and outward as the pressurized
air drives movement of the shroud-ring support 34. The flange
includes a radially inner flex-section 68, a radially outer
flex-section 70, and a mount section 72 as shown in FIG. 2. The
inner-flex section 68 is coupled to the panel 60 and extends
axially forward from the panel 60. The outer flex-section 70 is
coupled to the inner flex-section 68 and extends axially aft from
the inner flex-section 68. The inner and outer flex sections 68, 70
are arranged at an angle 74 relative to one another to provide the
generally u-shape of the flange 62. The angle 74 between the inner
flex-section 68 and the outer flex section 70 increases as the
diameter of the shroud-ring support 34 decreases. The angle 74
between the inner flex-section 68 and the outer flex section 70
decreases as the diameter of the shroud-ring support 34 increases.
The mount section 72 extends radially outward from the outer
flex-section 70 and couples to the outer case 44 to mount the
shroud-ring support 34 and the turbine shroud ring 26 to the outer
case 44.
In the illustrative embodiment, the passive blade tip clearance
control system 40 includes a plurality of inlet conduits 42 spaced
circumferentially around the central reference axis 11 as shown in
FIG. 3. Each of the inlet conduits 42 is a tubular member and
defines a bleed air passageway 48 that opens into the manifold 58.
The pressurized bleed air is conducted into the manifold 58,
through the gap 64 between the shroud-ring support 34 and the inner
panel 54 of the outer case 44, and then through one or more outlets
76 where the pressurized air is then sent to the low pressure
section 22 of the turbine 18. The outlet(s) 76 may be tubular
members similar to the inlet conduits 42 or may be defined by the
outer and/or the inner cases 44, 34 as will be described in greater
detail below.
As shown in FIG. 3, in the illustrative embodiment, the outlets 76
are spaced circumferentially around the reference axis 11. The
outlets 76 are tubular members defining passageways that extend
axially through the manifold 58 from the cavity 50 toward the low
pressure section 22 of the turbine 18. Each of the outlets 76 is
spaced apart circumferentially from neighboring inlet conduits 42
so that the pressurized air travels circumferentially from the
inlet conduits 42 toward the outlets 76 to deliver the pressurized
air over the entire circumferential area of the shroud-ring support
34. Additionally, the outlets 76 are spaced radially outward from
the inner panel 54 of the outer case 44 such that the pressurized
air travels radially outward over the inner panel 54 and then
through the outlets 76.
Turning again to FIG. 2, the high pressure section 20 of the
turbine 18 includes a first turbine blade stage 80, a second
turbine blade stage 82 axially aft of the first turbine blade stage
80, and a vane stage 84 axially between the first and second
turbine blade stages 80, 82. In the illustrative embodiment, the
passive blade tip clearance control system 40 is sized and located
to control a blade-tip clearance gap 36 radially between second
turbine blade stage 82 and the turbine shroud ring 26. As shown in
FIG. 2, the second turbine blade stage 82 generally falls within
the length 66 of the inner panel 54 of the outer case 44.
Pressurized air traveling through the gap 64 along the length 66
drives motion of the shroud-ring support 34 to move the turbine
shroud ring 26 radially inward or outward directly adjacent the
second turbine blade stage 82. As previously described, the length
66 of the inner panel 54 may be increased or decreased to target
additional and/or other turbine blade stages, such as first turbine
blade stage 80, and control the blade tip clearance of those
turbine blade stages. Additionally, the turbulators 90 facilitate
heat transfer in the gap 64.
Another embodiment of a passive blade tip clearance control system
240 in accordance with the present disclosure is shown in FIG. 4.
The passive blade tip clearance control system 240 is substantially
similar to the passive blade tip clearance control system 40 shown
in FIGS. 1-3 and described herein. Accordingly, similar reference
numbers in the 200 series indicate features that are common between
the passive blade tip clearance control system 40 and the passive
blade tip clearance control system 240. The description of the
passive blade tip clearance control system 40 is incorporated by
reference to apply to the passive blade tip clearance control
system 240, except in instances when it conflicts with the specific
description and the drawings of the passive blade tip clearance
control system 240.
The passive blade tip clearance control system 240 includes an
inlet conduit 242, an outer case 244 and a shroud-ring support 234,
or inner case 234. The outer case 244 and the shroud-ring support
234 define a cavity 250 radially therebetween. The inlet conduit
242 defines a bleed-air passageway 246 that is fluidly coupled
between the compressor 14 and the cavity 250 without interruption
from a valve or any other active blade tip clearance control
device. A cooling-air passageway 247 extends from the cavity 250 to
the low pressure section 22 of the turbine 18 and is also
uninterrupted from a valve or active blade tip clearance device. In
this way, the blade tip clearance control system 240 conducts
pressurized air from the compressor 14 into the cavity 250 to
control blade tip clearance in the high pressure section 20 and
then reuses the air downstream of the cavity 250 in the low
pressure section 22. Furthermore, the temperature and motion of the
shroud-ring support 234 is controlled based on the operating
conditions of the engine without active control of the pressurized
bleed air provided to the cavity.
The outer case 244 includes an outer panel 252, an inner panel 254
spaced apart radially from the outer panel 252, and an linking
segment 256 connecting the outer panel 252 to the inner panel 254
as shown in FIG. 4. The outer panel 252 and the inner panel 254
cooperate to define a radially-outer boundary of the cavity 250.
The inner panel 254 is spaced apart radially between both the outer
panel 252 and the shroud-ring support 234. The linking segment 256
extends from the outer panel 252 radially inward to the inner panel
254 at an angle to narrow the cavity 250 between the inner panel
254 and the turbine-shroud ring 234 to provide a gap 264 similarly
to the passive blade tip clearance control system 40 described
above.
The shroud-ring support 234, or inner case 234, includes a panel
260, a flange 262 on an axially forward end of the panel 260, and a
plurality of mounts 263 that extend radially downward from the
panel 260 as shown in FIG. 4. The panel 260 defines a
radially-inner boundary of the cavity 250. The flange 262 has a
generally U-shape cross section when viewed circumferentially and
extends radially outward into engagement with the outer case 244 to
provide the cavity 250. The flange 262 is configured to flex to
allow the shroud-ring support 234 to move radially inward and
outward as the pressurized air drives movement of the shroud-ring
support 234. The plurality of mounts 263 interface with
corresponding mounts 65 coupled to the turbine shroud ring 26 to
mount the turbine shroud ring 26 to the shroud-ring support 234. In
the illustrative embodiment, the mounts 263, 65 are generally
L-shaped when viewed circumferentially, however, in other
embodiments, the mounts may have any suitable shape. Additionally,
one or more fasteners may be used to mount the turbine shroud ring
26 to the shroud-ring support 234.
The bleed-air passageway 246 of the inlet conduit 242 has an outlet
248 that opens into the cavity 250 through the outer panel 252
axially forward from the inner panel 254 and the linking segment
256 as shown in FIG. 4. The pressurized air flows radially through
the outlet 248 and axially aft through the cavity 250 where it is
accelerated between the inner panel 254 and the shroud-ring support
234 along a length 266 of the inner panel 254. The length 266 of
the inner panel 254 corresponds to a portion of the shroud-ring
support 234 to encourage heat transfer between the pressurized air
and the shroud-ring support 234 in that area.
The inner panel 254 may be extended or retracted axially to adjust
the length 266 and, thus, the portion of the shroud-ring support
234 that is controlled. The size of the outer panel 252 or the
linking segment 256 may also be adjusted in view of the size of the
inner panel 254. Additionally, length 266 of the inner panel 254
may be increased or decreased based on the operating conditions of
the particular engine in which the passive blade tip clearance
control system is included. In one embodiment, an amount of radial
movement of the shroud-ring support 234 is related to the length
266 of the inner panel 254. For example, if more blade tip
clearance is required in a particular engine, the length 266 of the
inner panel 254 may be increased to provide greater heat transfer
across a larger portion of the shroud ring support 234 and, thus, a
larger change in diameter of the shroud ring support 234 relative
to the reference axis 11. However, it should be noted that any
suitable length 266 may be used depending on the particular
application.
The shroud-ring support 234 may further include a plurality of
turbulators 290 formed on the shroud ring support 234 as shown in
FIG. 4. The turbulators 290 are positioned in the gap 264 within
the length 266 of the inner panel 254 to increase the surface area
of the shroud ring support 234 in that area. The turbulators 290
increase heat transfer from the pressurized air to the shroud-ring
support 234 to help drive the shroud ring support 234 radially
inward and outward. However, in other embodiments, the turbulators
290 may extend along the shroud-ring support 234 a greater or
lesser distance than the length 266 of the inner panel 254.
In the illustrative embodiment, the passive blade tip clearance
control system 240 includes at least one outlet 276 down steam of
the gap 264 as shown in FIG. 4. The pressurized bleed air is
conducted through the gap 264 between the shroud-ring support 234
and the inner panel 254 of the outer case 244 and then through the
one or more outlets 276 where the pressurized air is then sent to
the low pressure section 22 of the turbine 18. In the illustrative
embodiment, the outlet(s) 276 are a void defined between aft ends
of the outer case 244 the shroud-ring support 234.
In the illustrative embodiment, the passive blade tip clearance
control system 240 is sized and located to control the blade-tip
clearance gap 36 radially between second turbine blade stage 82 and
the turbine shroud ring 26. The second turbine blade stage 82
generally falls within the length 266 of the inner panel 254 of the
outer case 244. Pressurized air traveling through the gap 264 along
the length 266 drives motion of the shroud-ring support 234 to move
the turbine shroud ring 26 radially inward or outward directly
outboard of the second turbine blade stage 82. As previously
described, the length 266 of the inner panel 254 may be increased
or decreased to target additional and/or other turbine blade
stages, such as first turbine blade stage 80, and control the blade
tip clearance of those turbine blade stages.
Another embodiment of a passive blade tip clearance control system
340 in accordance with the present disclosure is shown in FIG. 5.
The passive blade tip clearance control system 340 is substantially
similar to the passive blade tip clearance control system 40 shown
in FIGS. 1-3 and described herein. Accordingly, similar reference
numbers in the 300 series indicate features that are common between
the passive blade tip clearance control system 40 and the passive
blade tip clearance control system 340. The description of the
passive blade tip clearance control system 40 is incorporated by
reference to apply to the passive blade tip clearance control
system 340, except in instances when it conflicts with the specific
description and the drawings of the passive blade tip clearance
control system 340.
The passive blade tip clearance control system 340 includes an
inlet conduit 342, an outer case 344 and a shroud-ring support 334,
or inner case 334. The outer case 344 and the shroud-ring support
334 define a cavity 350 radially therebetween. The inlet conduit
342 defines a bleed-air passageway 346 that is fluidly coupled
between the compressor 14 and the cavity 350 without interruption
from a valve or any other active blade tip clearance control
device. A cooling-air passageway 347 extends from the cavity 350 to
the low pressure section 22 of the turbine 18 and is also
uninterrupted from a valve or active blade tip clearance device. In
this way, the blade tip clearance control system 340 conducts
pressurized air from the compressor 14 into the cavity 350 to
control blade tip clearance in the high pressure section 20 and
then reuses the air downstream of the cavity 350 in the low
pressure section 22. Furthermore, the temperature and motion of the
shroud-ring support 334 is controlled based on the operating
conditions of the engine without active control of the pressurized
bleed air provided to the cavity.
The outer case 344 includes an outer panel 352 and an inner panel
354 as shown in FIG. 5. The outer panel 352 and the inner panel 354
cooperate to define a radially-outer boundary of the cavity 350.
The inner panel 354 is coupled to the outer panel 352 and is spaced
apart radially outward of the shroud-ring support 334. The inner
panel 354 is arranged to narrow the cavity 350 between the inner
panel 354 and the turbine-shroud ring 334 to provide a gap 364
similarly to the passive blade tip clearance control system 40
described above.
The shroud-ring support 334, or inner case 334, includes a panel
360, a flange 362 on an axially forward end of the panel 360, and a
plurality of mounts 363 that extends radially downward from the
panel 360 as shown in FIG. 5. The panel 360 defines a
radially-inner boundary of the cavity 350. The flange 362 has a
generally U-shape cross section when viewed circumferentially and
extends radially outward into engagement with the outer case 344 to
provide the cavity 350. The flange 362 is configured to flex to
allow the shroud-ring support 334 to move radially inward and
outward as the pressurized air drives movement of the shroud-ring
support 334. The plurality of mounts 363 interface with
corresponding mounts 65 coupled to the turbine shroud ring 26 to
mount the turbine shroud ring 26 to the shroud-ring support 334. In
the illustrative embodiment, the mounts 363, 65 are generally
L-shaped when viewed circumferentially, however, in other
embodiments, the mounts may have any suitable shape. Additionally,
one or more fasteners may be used to mount the turbine shroud ring
26 to the shroud-ring support 334.
The bleed-air passageway 346 of the inlet conduit 342 has an outlet
348 that opens into the cavity 350 through the outer panel 352
axially forward from the inner panel 354 as shown in FIG. 5. The
pressurized air flows radially through the outlet 348 and axially
aft through the cavity 350 where it is accelerated between the
inner panel 354 and the shroud-ring support 334 along a length 366
of the inner panel 354. The length 366 of the inner panel 354
corresponds to a portion of the shroud-ring support 334 to
encourage heat transfer between the pressurized air and the
shroud-ring support 334 in that area.
In the illustrative embodiment, the inner panel 354 is a separate
component that is mounted to a radially-inner surface of the outer
panel 352 as shown in FIG. 5. The inner panel 354 includes an
axially-forward segment 394, an axially-aft segment 396, and a
middle segment 398 extending between the axially-forward segment
394 and the axially-aft segment 396. The axially-forward segment
394 extends radially inward at an angle relative to the reference
axis 11 to the middle segment 398. The axially-aft segment 396
extends radially outward at an angle relative to the reference axis
11 from the middle segment 398 to the outer panel 352. In this way,
the axially-forward segment 394 and the axially-aft segment 396
cooperate to locate the middle segment 398 in spaced-apart relation
to the outer panel 352 and the shroud-ring support 334 and to
narrow the cavity 350 along the length 366 of the inner panel 354.
The middle segment 398 is generally parallel with the shroud-ring
support 334 to provide a constant gap 364 between the middle
segment 398 and the shroud-ring support 334. The segments 394, 396
may be coupled to the outer panel 352 by welding, brazing,
mechanical fasteners or any other suitable means of mounting the
inner panel 354 to the outer panel 352.
The inner panel 354 may be extended or retracted axially to adjust
the length 366 and, thus, the portion of the shroud-ring support
334 that is controlled. More particularly, the length of the middle
segment 398 is increased or decreased in some embodiments.
Additionally, length 366 of the inner panel 354 may be increased or
decreased based on the operating conditions of the particular
engine in which the passive blade tip clearance control system is
included. In one embodiment, an amount of radial movement of the
shroud-ring support 334 is related to the length 366 of the inner
panel 354. For example, if more blade tip clearance is required in
a particular engine, the length 366 of the inner panel 354 may be
increased to provide greater heat transfer across a larger portion
of the shroud ring support 334 and, thus, a larger change in
diameter of the shroud ring support 334 relative to the reference
axis 11. However, it should be noted that any suitable length 366
may be used depending on the particular application.
The shroud-ring support 334 may further include a plurality of
turbulators 390 formed on the shroud ring support 334 as shown in
FIG. 5. The turbulators 390 are positioned in the gap 364 and
within the length 366 of the inner panel 354 to increase the
surface area of the shroud ring support 334 in that area. The
turbulators 390 increase heat transfer from the pressurized air to
the shroud-ring support 334 to help drive the shroud ring support
334 radially inward and outward. However, in other embodiments, the
turbulators 390 may extend along the shroud-ring support 334 a
greater or lesser distance than the length 366 of the inner panel
354.
In the illustrative embodiment, the passive blade tip clearance
control system 340 includes at least one outlet 376 downstream of
the gap 364 as shown in FIG. 5. The pressurized bleed air is
conducted through the gap 364 between the shroud-ring support 334
and the inner panel 354 of the outer case 344 and then through the
one or more outlets 376 where the pressurized air is then sent to
the low pressure section 22 of the turbine 18. In the illustrative
embodiment, the outlet(s) 376 are a void defined between aft ends
of the outer case 344 the shroud-ring support 334.
In the illustrative embodiment, the passive blade tip clearance
control system 340 is sized and located to control the blade-tip
clearance gap 36 radially between second turbine blade stage 82 and
the turbine shroud ring 26. The second turbine blade stage 82
generally falls within the length 366 of the inner panel 354 of the
outer case 344. Pressurized air traveling through the gap 364 along
the length 366 drives motion of the shroud-ring support 334 to move
the turbine shroud ring 26 radially inward or outward directly
outboard of the second turbine blade stage 82. As previously
described, the length 366 of the inner panel 254 may be increased
or decreased to target additional and/or other turbine blade
stages, such as first turbine blade stage 80, and control the blade
tip clearance of those turbine blade stages.
The present disclosure relates to a passive tip clearance control
system for either high or low pressure turbines. The system may
include concentric inner 34 and outer turbine cases 44. A flange 62
with a flexible connecting section may join the forward end of the
inner case 34 to the forward end of the outer case 44. This
arrangement may allow the cases to move independently from one
another, and may create a cavity 50 between the cases in which air
can be circulated to control tip clearance 36. Blade track hangers
and blade tracks (collectively 26) are attached to the inner
diameter of the inner case 34.
In some embodiments, a manifold 58 mounted aft of the second stage
turbine blade track may receive air piped externally from the
compressor 14. The manifold accelerates and directs the air over
the outer surface of the inner case adjacent to the second stage
blade track. The outer surface of the inner case may have
turbulators 90, fins, pins, or other means of increasing the rate
of heat transfer with the passing air. The arrangement of these
features may be adjusted to achieve the desired response of the
case and blade track relative to the rotor and blade tip. The air
continues over the forward section of the inner case over the first
stage blade track, then turns to flow aft against the inner surface
of the outer case. Separate passages 76 in the manifold direct the
air back into the circuit to be used downstream in the turbine of
the engine.
In some embodiments, air piped from the compressor is fed through
the outer case 244, 344. A cavity between the inner and outer cases
may function as the manifold. A duct 64, 264, 364 is used to
accelerate and direct the air aft over the outer surface of the
inner case where it is radially aligned with the second stage blade
track.
In some embodiments, the passive blade tip clearance control system
may heat the case during the initial part of the mission when the
thermal and mechanical growth of the rotor (or turbine wheel 24)
tends to outpace the thermal growth of the case, and then cool the
case during the cruise portion of the mission when the case tends
to thermally expand away from the rotor. This system may open the
tip clearance 36 during the initial part of the mission to avoid
contact between the blades and blade track (contact may result in a
permanent increase in tip clearance), and tightens the tip
clearance 36 during the cruise portion of the mission to improve
efficiency. The passive blade tip clearance control system may
provide these effects in a simple and robust way without the need
of active valves, mechanical actuators, or complex control systems
that use air impingement, valves, mechanical actuation, or some
combination of these to control tip clearance.
While the disclosure has been illustrated and described in detail
in the foregoing drawings and description, the same is to be
considered as exemplary and not restrictive in character, it being
understood that only illustrative embodiments thereof have been
shown and described and that all changes and modifications that
come within the spirit of the disclosure are desired to be
protected.
* * * * *